The present invention relates generally to gas turbine engines, and, more specifically, to turbine nozzles therein.
In a gas turbine engine, air is pressurized in a compressor and mixed with fuel and ignited in a combustor for generating hot combustion gases. The gases are discharged through a first stage high pressure turbine nozzle having stator vanes which direct the gases toward a row of turbine rotor blades extending radially outwardly from a supporting disk.
The turbine blades extract energy from the combustion gases and power the compressor. The gases are then channeled to a low pressure turbine typically having several stages of nozzle vanes and rotor blades which extract additional energy from the gases for producing output work such as powering a fan in a turbofan aircraft engine embodiment.
Since the high pressure turbine nozzle firstly receives the combustion gases from the combustor, it must be cooled for enjoying a suitable useful life. A typical turbine nozzle includes a row of airfoil vanes circumferentially spaced apart from each other and extending radially in span between outer and inner annular bands. The vanes are hollow for receiving therein a portion of compressor discharge air used for cooling the individual vanes.
Internal cooling channels are defined in each vane by corresponding radially extending ribs or partitions which integrally join together the circumferentially opposite pressure and suction sides of the vane. The inner surfaces of the vanes may include short turbulators which trip the cooling air flowing thereover during operation for enhancing heat transfer cooling therefrom.
In order to protect the external surface of the vanes from the hot combustion gases flowing thereover, various radial rows of film cooling holes are provided through the pressure and suction sides of the vane. Since the leading edge of the vane first receives the hot combustion gases, it typically includes several rows of film cooling holes in a showerhead configuration. The air discharged from the film cooling holes produces a boundary layer of cooling air along the external surface of the vane which is re-energized with additional cooling air from row-to-row. The film cooling air provides a barrier protecting the metal of the vane from the hot combustion gases during operation.
A typical vane airfoil increases in thickness aft of the leading edge to a maximum thickness typically within the first third of the chord length, and then tapers and narrows in thickness to a relatively thin trailing edge. As the vane thins near the trailing edge, the ability to cool the trailing edge region of the vane becomes more difficult. The trailing edge is thusly another region of the vane which experiences relatively high temperature during operation.
The trailing edge is typically cooled by a row of trailing edge discharge holes which provide internal convection cooling thereof. And, one or more rows of additional film cooling holes may be provided along the pressure sidewall for protecting the pressure sidewall and developing a cooling air film which extends downstream to the trailing edge for the additional protection thereof.
Furthermore, the suction sidewall may also include several rows of film cooling gill holes between the leading edge and the maximum thickness region which develop cooling air films for protecting the suction sidewall, and which flow to the trailing edge for the additional protection thereof.
Since the combustion gases flow with different velocities over the pressure and suction sidewalls of the vane, the various regions of the vane from leading to trailing edge are subject to different amounts of heating therefrom, and correspondingly require different amounts of cooling. Since any air diverted from the combustor for cooling the nozzle vanes decreases overall engine efficiency, the amount thereof should be minimized while obtaining a suitable useful life for the nozzle vanes.
The varying heating effect of the combustion gases, and the varying cooling effect of the cooling air further complicate vane design since temperature gradients are created. Temperature gradients cause differential expansion and contraction of the vane material, which in turn causes thermally induced strain and stress which affects the low cycle fatigue life of the vane during operation.
For example, partitions or ribs extend between the pressure and suction side of the vane to define corresponding cooling channels therein and are inherently relatively cold since they are protected inside the vane and cooled by the air channeled therealong. The ribs are relatively cold when compared with the relatively hot pressure and suction sidewalls of the vane, and a considerable temperature gradient is created therebetween. Furthermore, temperature gradients are also effected between the leading and trailing edges of the vane in different amounts along the pressure and suction sides.
Accordingly, the prior art is crowded with various configurations for cooling turbine nozzle vanes with different complexity and different degrees of effectiveness, and with different useful lives.
For example, General Electric Company has manufactured and sold one turbofan aircraft gas turbine engine designated as the CF34 model which has enjoyed decades of commercial success and use. The high pressure turbine nozzle of this engine includes film cooled vanes having a significant useful life. Decades of commercial use of this engine has provided thousands of hours of field experience for evaluating durability and life of the turbine nozzles therein.
Such field experience in conjunction with extensive analysis of this nozzle design may now be used for improving the durability and life of the turbine nozzle without increasing the amount of cooling air required therefor.
Accordingly, it is desired to provide an improved turbine nozzle based on extensive field experience and analysis having improved durability without requiring additional cooling airflow.